The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the gases in a high pressure turbine (HPT) which powers the compressor, and additional energy is extracted in a low pressure turbine (LPT) which powers an upstream fan in an exemplary turbofan aircraft engine application, or drives an output drive shaft in marine and industrial applications.
The HPT and LPT include corresponding turbine nozzles which are heated by the hot combustion gases during operation. The first stage nozzles in the HPT and LPT each includes a row of hollow stator vanes integrally joined at opposite ends thereof to radially inner and outer bands. The stator vanes have identical airfoil configurations for channeling the combustion gases to a downstream row of turbine rotor blades extending radially outwardly from the perimeter of a supporting rotor disk.
The stator vanes also include identical cooling circuits therein in which corresponding partitions extend between the opposite pressure and suction sides of the airfoils to define several flow channels extending radially in span between the two bands for circulating in the airfoil a portion of pressurized cooling air bled from the compressor during operation.
The cooling circuits may have various conventional configurations and commonly include various rows of film cooling holes extending through the opposite pressure and suction sides of each airfoil for discharging the spent cooling air in corresponding thin films of cooling air that create thermally insulating blankets of air for externally protecting the vanes from the hot combustion gases.
The combustion gases first reach the stator vanes along the leading edges thereof which are typically protected by corresponding rows of film cooling or gill holes distributed along the span of each airfoil. The combustion gases then flow through corresponding nozzle channels between adjacent vanes and leave the nozzle along the trailing edges of the vanes. The thin trailing edges typically include a row of film cooling outlet slots that discharge another portion of the spent cooling air from the internal cooling circuits.
Since the nozzle vanes have specific aerodynamic airfoil configurations for use in extracting energy in the HPT and LPT, the velocity and pressure distributions of the combustion gases over the differently configured pressure and suction sides of each vane are different, and correspondingly create different heat loads on the vanes. Furthermore, each vane is fixedly mounted at its opposite ends to the outer and inner bands which are also subject to the heat loads of the combustion gases.
Since the nozzle is an annular structure, the heat loads from the combustion gases cause the nozzle to expand in diameter as it is heated, and to correspondingly contract in diameter as the heat, and corresponding operating temperature, are reduced.
This expansion and contraction of the annular turbine nozzle due to the change in heat loads from the combustion gases creates significant thermal stresses in both the individual vanes and supporting bands. Since the gas turbine engine operates in repeating cycles over its expected lifetime, the turbine nozzle is subject to heating and cooling cycles which introduce low cycle fatigue (LCF).
The life of the turbine nozzle is typically limited by the accumulating cycles of LCF experienced by the nozzle. The LCF life of the nozzle is typically limited by any one location in the nozzle that experiences the most accumulated fatigue from the LCF cycles which could eventually lead to a corresponding reduction in nozzle strength and the introduction of undesirable crack damage in the nozzle.
However, the LCF limited location in the turbine nozzle is a function of the specific operating cycle of the engine, and of the specific design of the nozzle itself including its specific cooling configuration.
The LCF life of the typical turbine nozzle is correspondingly increased by circumferentially dividing the annular nozzle into small segments typically including one or two nozzle vanes in corresponding segments of the outer and inner bands. Segmenting the annular nozzle interrupts the hoop continuity thereof and reduces the magnitude of thermal stresses therein.
However, segmenting the annular turbine nozzle correspondingly requires suitable seals between those segments which increase the complexity of the nozzle, and may reduce its overall efficiency.
A single vane nozzle segment fully uncouples the circumferential continuity of the annular nozzle, and allows the individual nozzle vane to freely expand and contract with its corresponding band segments.
In a two vane nozzle segment, or doublet, the individual vanes are no longer free to expand and contract alone, but are subject to the expansion and contraction of the second vane and its integral connection to the common band segments.
And, a three vane nozzle segment, or triplet, further increases thermal restraint since any one of the three vanes is integrally interconnected with the other two vanes by the corresponding band segments.
In particular, thermal expansion of the three vane segment tends to straighten the circumferential curvature or arc of the outer band which introduces tensile loads in the two outboard or end vanes, while correspondingly introducing compression loads in the middle or mid vane.
Since the vanes themselves, their cooling configurations, and the mounting band segments are typically identical from segment to segment, the different thermal loads experienced in the segments introduce different thermal stresses in the vanes and bands which adversely affect the LCF life of the nozzle. Since the middle vane in the three vane nozzle segment is structurally trapped by the two end vanes and the common supporting bands, the LCF life limiting location may be found in the middle vane.
The complexity of the modern gas turbine engine turbine nozzles and their hostile operating environment therefore require a balance of engine performance or efficiency and LCF life.
Accordingly, it is desired to provide a turbine nozzle having an improved configuration for increasing both aerodynamic efficiency and LCF life thereof.